Oblique blended wing body aircraft

ABSTRACT

An oblique wing aircraft designed for reduced surface area to volume ratio. The aircraft has an oblique wing comprising a forward swept wing segment and an aft swept wing segment. A center oblique airfoil section connects the forward and aft swept wing segments. The center oblique airfoil section has a larger chord near its centerline than the chords of either of the forward or aft swept wing segments. The chord of the center oblique airfoil section tapers down more rapidly than the forward or aft wing segments as the center oblique airfoil section extends outboard toward the forward and aft swept wings. Preferably, the aircraft is an all-wing aircraft.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.13/784,438, filed Mar. 4, 2013, which is continuation of U.S. NationalStage application Ser. No. 12/675,165, filed Feb. 25, 2010, which is a35 U.S.C. §371 of International Application No. PCT/US08/074795, filedAug. 29, 2008, which in turn claims priority to U.S. Provisional PatentApplication Ser. No. 60/935,758, filed Aug. 29, 2007, all which arehereby incorporated by reference.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH

Not Applicable.

TECHNICAL FIELD

The present invention relates to flying objects designed to fly fastenough that compressibility drag becomes significant.

BACKGROUND ART

In the past people have proposed blended wing body aircraft like the B-2bomber and people have also proposed all flying Oblique Wings as shownin FIG. 1 with an elliptical or near elliptical planform.

Blended Wing Body aircraft like the B-2 achieve lower drag than a pureflying wing by minimizing the surface area exposed to the airflow. Theydo this by having a center body that is as close as practical tocircular in planform but usually with a pointed nose on the front toreduce compressibility drag and with wings attached to the sides toincrease the wingspan for reduced induced drag which is drag due tocreating lift. A wing with a circular planform has the least amount ofsurface area to internal volume for the same reason that a circle hasthe smallest circumference to the enclosed area or a sphere has thelargest volume to surface area. The Blended Wing Body aircraft also canhave inherent pitch stability at a farther aft center of gravity due tothe aft swept wings that can act like horizontal tail surfaces. Furtherbackground of blended wing body aircraft is given in R. H. Liebeck,“Design of the Blended-Wing-Body Subsonic Transport,” 2002 WrightBrothers Lecture, American Institute of Aeronautics and Astronautics,AIAA-2002-0002, reprinted in Journal Of Aircraft, Vol. 41, No. 1,January-February 2004, pp. 10-25, hereby incorporated by reference.

Oblique flying wing aircraft that have been proposed in the past wereelliptical or near elliptical wings that flew at different obliqueangles to trade off compressibility and induced drag at different machnumbers like that shown in planform in FIG. 1. The design shown in FIG.1 has remained relatively unchanged since it was proposed by R. T. Jonesin the 1950's. The history of oblique wing research is found in M.Hirschberg, D. Hart, and T. Beutner, “A Summary of a Half-Century ofOblique Wing Research,” 45th AIAA Aerospace Sciences Meeting andExhibit, AIAA Paper 2007-150, January 2007, hereby incorporated byreference.

At low speed the aircraft could fly in a low speed direction 2 close toa zero sweep angle for minimum induced drag which is the drag due tolift. At high speed, compressibility drag becomes more important andeventually dominant. Compressibility drag due to lift andcompressibility drag due to volume can however be reduced by spreadingthe lift and volume farther in the direction of flight. Thus as theaircraft flew faster and faster the wing was swept to a higher andhigher sweep angle to trade off the optimum induced versuscompressibility drag characteristics. The component of air velocityperpendicular to the wing could remain subsonic effectively making thewing and air interact very similar to a wing flying subsonically.Engines 6 were generally envisioned to be mounted in rotating pods onthe bottom of the wing. The small chord length and limited thickness ofthe wing made integrating the engine into the wing more difficult and inorder to have an aircraft with a thick enough wing that passengers couldstand up in a cabin the aircraft had to be very large carryingapproximately six hundred passengers. The largest circle possible 5 isshown drawn over (inscribed in) the planform of the aircraft shown inFIG. 1. As may be seen, it is a small circle encompassing only a smallpercentage of the planform area of the aircraft. From this we candetermine that this aircraft has a large amount of surface area tointernal volume ratio and as such will have a lot of skin friction dragin both high and low speed configurations. Also because the circle issmall and there is a finite limit to the thickness to chord length ofthe airfoil used on this flying wing, we know the thickness of thevehicle will not be very large making packaging of the vehicle moredifficult or requiring the vehicle to be larger than might be desirablesuch as to incorporate a cabin for passengers or other components.

In the past people have also proposed oblique wing aircraft that hadconventional fuselages as well. Problems occurred due to the interactionbetween the wing and fuselage, and the high compressibility drag due tovolume of the fuselage caused most designers to look to all-wingconfigurations.

Oblique flying wing aircraft have more surface area to volume than aBlended Wing Body aircraft like the B-2 stealth bomber and they alsoneed the center of gravity very far forward or they are unstable andhard to control and generally have to be provided with an advancedartificial stabilization system.

SUMMARY OF THE INVENTION

The present invention combines the benefits of a Blended Wing Bodyaircraft and an all flying Oblique Wing Aircraft. The invention combinesthe low wetted area and natural stability benefits of a blended wingbody aircraft with the variable sweep and low compressibility dragbenefits of an Oblique flying wing.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a plan view of a traditional Oblique All-Wing Aircraft (priorart).

FIG. 2 is a plan view of one embodiment of the present invention.

FIG. 3 is a plan view of another embodiment of the present invention.

FIG. 4 is a view looking aft at the embodiment of the invention shown in

FIG. 2 in low speed configuration perpendicular to the leading edge 9 ofthe forward swept wing 27 and center body 25.

FIG. 5 is a cross-sectional cut from FIG. 2 looking inboard from theleft wing parallel to the leading edge 9, as indicated by line 5-5 ofFIG. 2.

BEST MODE FOR CARRYING OUT THE INVENTION

Referring to FIG. 2, a center airfoil section, called herein the “centerbody” 25, is designed by starting off with a circle 7 as shown in dashedlines at the center of the aircraft planform. Tapering transition areas15 and 13 are added on either side of the circle to form left and rightsides of the center body 25 and a forward swept wing 27 and an aft sweptwing 29 are then attached to the center body 25 as shown in FIG. 2 planview. The center body 25 is itself a wing segment with airfoilcross-sections. The forward swept wing 27 preferably has a leading edge9 that is an extension of the leading edge of the center body's forwardswept transition area 15. The aft swept wing 29 has its aerodynamiccenter behind the pitch axis 39 so it is helpful in stabilizing thevehicle in pitch about the pitch axis 39. The pitch axis 39 for thisconfiguration is defined parallel to the quarter chord of the forwardswept wing but passing through the vehicle center of gravity as shown inFIG. 2. Preferably the aft swept wing aerodynamic center is fartherbehind the pitch axis 39 than the forward swept wing aerodynamic centeris in front of this pitch axis. In order to balance the aircraft inpitch and roll with minimal control surface deflections for lowest drag,the aft swept wing 29 for the configuration shown in FIG. 2 preferablyhas a positive dihedral and a lower incidence relative to the forwardswept wing 27. For yaw trim at high sweep angles, a vertical fin 23 canbe located at the end of the aft swept wing 29. The current inventionhas less of a problem with the oblique wing wanting to yaw to lowersweep angles than with a traditional oblique wing aircraft as will bedescribed later. However if, for the particular design chosen, theaircraft still wants to rotate to lower sweep angles, then more verticalfin 23 area should be located above the wing than below. This fin areaabove the wing generates an inboard force (in this case to the right) tokeep the aircraft from naturally wanting to rotate back to a lower sweepangle. By having the fin located more above the wing, it also acts as awinglet to reduce induced drag since an upward facing winglet willnaturally produce an inboard force and a fin located above the wingpushing inboard will naturally tend to act as a winglet by preventinghigh pressure air below the wing from moving around the tip of the wingand onto the top surface of the wing. If the aft swept wing 30 is sweptfar aft such as shown in the configuration shown in FIG. 3, the aircraftcan be more inherently stable under some circumstances or the center ofgravity can be farther aft for the same amount of natural stability.However, with the aft swept wing oriented as shown in FIG. 2, thevehicle has a higher aspect ratio and has the potential for lowerinduced drag at the same forward swept wing sweep angles. At low speeds,both forward and aft facing wings of the configuration shown in FIG. 2can have zero sweep making them more adaptable to using laminar flowairfoils.

Another way to think of the current invention is a low aspect ratiooblique all wing aircraft to achieve a low surface area to volume ratiobut with the addition of wing tip extensions to achieve more inherentpitch stability and achieve the necessary wingspan for low induced dragwhich is drag due to generating lift. These shorter chord wing tipextensions are designed to generate much more lift per square foot ofplanform area than the center oblique wing section to compensate fortheir smaller chord. They do this by operating at a higher liftco-efficient usually from increased angle of attack due to either 1)their position in the upwash field downstream of the other liftingsurfaces or 2) increased pitch angle due to increased incidence or 3)increased pitch from the dihedral on the forward swept wing. The wingtipextensions also usually generate more lift co-efficient because thecenterbody usually will use a reflex airfoil which has a lower liftco-efficient but provides a desirable nose up pitching moment. The wingtip extensions can also generate a higher lift co-efficient withdeflected trailing edge flaps or other lift augmenting devices thoughthis is generally not the preferred approach. The wingtip extensionsshould be designed to provide twice (and preferably three times) or morethe average lift per square foot of wing area as the center oblique wingsection but lower numbers such as 30% to 60% more lift per square footmay make sense for some designs wanting a very high level of maneuvercapability before stalling the wingtip extensions.

A. How to Lay Out the Planform of the Current Invention

FIG. 2 shows a plan view of one embodiment of the current invention.

The largest possible circle 7 has been drawn over (inscribed in) theaircraft planform. Relative to previous Oblique Wing aircraft as shownin FIG. 1, a much larger percentage of the planform area fits insidethis circle resulting in a much lower ratio of surface area to internalvolume for this configuration and thus lower skin friction drag. As aresult designing an optimum configuration of the current invention canstart by drawing a circle. On either side of the circle we then drawlines tangential to the circle and rapidly tapering down going forwardand to the right, and aft and to the left (or vice versa for a mirrorimage layout) in order to create the transition areas 13 and 15 on leftand right sides of the center body 25. Finally, the wings 29 or 30 and27 are attached to the left and right sides of the center body 25. Thefaster the chord of the wing tapers down in transition areas 13 and 15on either side of the circle 7, the lower, and better, the surface areato volume ratio will be. However the slower it tapers down generally thelower the compressibility drag associated with volume can be. As aresult this is a design trade. If the top speed of the aircraft is veryhigh then the wing chord taper in the transition areas 13 and 15 willideally be more gradual, and the slower the top speed of the aircraft,the faster the chord will taper down. Even for a high speed vehicle thearea inside the center circle should be more than 25% of the totalplanform area and typically 33% or more of the planform area. For alower speed vehicle it should be more than 40% and preferably betterthan 50% of the total planform area. Above 55 to 60% is approaching anupper limit to a practical design that is trying to minimizecompressibility drag. The rate of taper up or taper down of chord lengthis defined by the difference in sweep angles between the leading andtrailing edge surfaces (9 versus 19 and 17 versus 21) in plan view. Inthe configurations shown the airfoil sections chosen for the aircraft todescribe the aircraft outer mold line are laid out perpendicular to thepitch axis 39. If the vehicle is used for military purposes and radarstealth is a consideration then the leading edges like 9 and 31 can beparallel to trailing edges like 21, and leading edge 54 can be parallelto trailing edge 19 in order to reduce the number of radar spikesincorporated in the design.

The faster the vehicle is designed to fly, the higher the sweep angle isdesired on the leading edge 9 of the forward swept wing 27 and itstransition area 15 relative to the direction of flight. However,generally the sweep angle on the leading edge 17 of the transition area13 should remain less than 90 degrees so that it stays a leading edgeboth in high and low speed flight. Likewise, the trailing edge 19 of theforward swept transition area 15 should be swept less than 90 degrees sothat it stays a trailing edge in both high and low speed flight angles.As a result, the taper down angle or angle between the leading 9 andtrailing 19 edges of the transition area 15 generally should be lessthan 90 degrees, and preferably less than 85 degrees, minus the desiredmaximum sweep angle of the leading edge 9 for the forward swepttransition area 15 relative to the direction of flight. Similarly, thetaper down angle or angle between the leading 17 and trailing 21 edgesof the aft swept transition area 13 should be less than 90 degrees, andpreferably less than 85 degrees, minus the desired maximum sweep anglerelative to the direction of flight of the trailing edge 21 for the aftswept transition area 13.

FIGS. 2 and 3 show two different sweep angles for alternative aft sweptwing configurations 29 and 30. Aft swept wing configuration 29 shown inFIG. 2 has the same or close to the same sweep angle as forward sweptwing 27 and thus both wings can be swept just behind the mach cone sothe air and wing is interacting very similar to how they would interactin subsonic flight with little or no shock waves, and yet achieve thelargest aspect ratio for lowest induced drag. Aspect ratio is thewingspan squared divided by the planform area of the wing. Theconfiguration of FIG. 2 probably represents the lowest desirable sweepangle for the aft swept wing.

The alternate aft swept wing configuration 30 shown in FIG. 3 however isgenerally lower risk from a pitch stability and control standpoint sincethe vehicle is generally more stable in pitch because the aerodynamiccenter of pressure of the aft swept wing 30 is farther aft andpotentially the center of gravity of the vehicle can be farther back forthe same level of inherent and artificial stability. Aft swept wing 30,as shown in FIG. 3, also places elevon control surfaces 60 and 61farther aft behind the pitch axis 39 so that they have a larger momentarm for providing pitch control for the vehicle. In FIG. 3, the hingeline 63 for elevon 61 is shown at a lower sweep angle than the aft sweptwing 30 in order to make the elevon 61 more effective aerodynamically.As can be seen, the hinge line 63 is still swept more relative to thehigh speed direction of flight 3 than the forward swept wing 27A so thatit is still swept behind the mach cone.

More highly swept wing 30 as shown in FIG. 3 will generally be easier todesign to handle the widely varying upwash angles generated by thecenter body 25 and forward swept wing 27A without stalling. However, thelift curve slope or change of lift co-efficient as a function of angleof attack will be steeper for aft swept wing 29 than it will for wing 30because of the lower sweep angle which will tend to reduce the pitchstability differences between the two configurations as long as wing 29is kept from stalling. If a vertical fin 23 is used, then aft swept wingconfiguration 30 also provides more vertical tail volume for the samesize vertical fin 23. Aft swept wings with sweep angles different fromthat of wings 29 and 30 are of course also possible but will generallybe between these two configurations. The increased sweep angle for theleading edge 18 of the aft swept wing 30 relative to the leading edge 9of the forward swept wing 27 should be less than 80 degrees, andpreferably less than 75 degrees, minus the maximum design sweep angle inhigh speed flight of the leading edge 9 of the forward swept wing 27.

B. Achieving Low Compressibility Drag Characteristics

The current invention achieves low compressibility drag which is thedrag associated with going close to or over the speed of sound(“transonic speed”) where the air acts like a compressible gas.Compressibility drag is kept low by having the airfoils swept in onedirection from one end of the vehicle to the other to keep the isobars,or lines of constant pressure, swept in one direction similar toprevious oblique wings. However this invention differs from previousoblique wings in that there is a more rapid buildup in thickness, chordand volume near the vehicle centerline. Relative to the forward sweptwing 27, and unlike previous oblique wings, the sweep angle of thequarter chord line 1 (shown in dotted line and shows the points onequarter of the way from the leading edge to the trailing edge of theairfoil) and the half chord line (not shown) preferably increases in thetransition areas 13 and 15 on either side of the centerline of thecenter body 25 where the chord and thickness of the airfoil is morerapidly tapering up or down. The increased sweep angle of the quarterand half chord lines tends to compensate for the increasedcompressibility drag due to volume that might otherwise be associatedwith this more rapid tapering up or tapering down of volume. Theincreased sweep gives the air more time to move out of the way of theincreasing cross- sectional area. In the preferred embodiments as shownin FIGS. 2 and 3, the quarter chord is more swept in the forward swepttransition area 15 because the trailing edge 19 is heavily swept aftwhile the leading edge 9 maintains the same sweep angle as the forwardswept wing 27 and 27A. Also the quarter chord is more swept in the aftswept transition area 13 because the leading edge 17 is heavily sweptaft while the trailing edge 21 stays at a modest sweep angle close tothat of the leading edge 9 of the forward swept wing 27 or 27A. Evengreater chord sweep angles could be achieved for example by increasingthe sweep of both leading 9 and trailing edges 19 in the forwardtransition area 15. Although not shown, a modified half chord line(showing a line created by points half way between the forward and aftedges of the airfoil as laid out perpendicular to this half chord linedrawn on the forward swept wing 27) could also be drawn on FIGS. 2 and 3and would have a similar appearance to the quarter chord line but wouldbetter represent the average shape between the leading and trailingedges of the vehicle. For the configurations shown in FIGS. 2 and 3, thehalf chord line is swept just over 20 degrees more in the forward 15 andaft 13 transition areas than in the forward swept wing 27. Thisincreased sweep angle for a high speed aircraft should be over 10degrees and preferably over 15 degrees. For a lower speed vehicle theincreased sweep angle tends to actually be higher and is made possibleby the fact that the forward swept wing doesn't operate to as high asweep angle. For a lower speed vehicle the increased sweep should beover 20 degrees and preferably over 25 degrees, with over 30 degreesapproaching a practical limit.

To reduce compressibility drag further the vehicle is preferably arearuled. A NACA researcher named Dr. Richard Whitcomb discovered that thewave drag is related to the second-derivative (or curvature) of thevolume distribution of the vehicle. The lowest wave drag occurs with aSears-Haack area distribution where the curvature of the volumedistribution is minimized. Although area ruling is less critical with anOblique wing it is still beneficial. Area ruling by adjusting theairfoil thicknesses makes it possible to retain the low radar crosssection design associated with only two sets of parallel lines todescribe the outline of the vehicle in plan view as shown in FIGS. 2 and3 that otherwise might cause larger curvature of the volume distributionthan desired at certain areas like at the connections of the wings 27and 29 to the transition areas 15 and 13.

C. How to Establish the Wing and Center Body Incidence

This aircraft is very unusual because of the highly unsymmetrical natureof the vehicle. As a result, the left and right sides of the vehiclescan be dramatically different. To establish the wing incidence oneshould start off with the aircraft in its low speed, low sweeporientation. For the configuration shown in FIG. 2, there is no need forthe vehicle to rotate past the point where both wings 27 and 29 are atright angles to the direction of flight. As a result, the quarter chordline 1 of the forward swept wing 27, even when the aircraft is at itslowest sweep angle, is forward of the chord line 1 of the aft swept wing29 and center body 25. As a result, the center body 25 generally will beflying in the upwash from the forward swept wing 27. Also the aft sweptwing 29 or 30 will be flying in the upwash from both the center body 25and the forward swept wing 27/27A. It is desirable to have an ellipticaldistribution of lift across the wingspan of the aircraft with thetrailing edge flaps as close to neutral as possible for minimum drag. Toachieve this the forward swept wing 27, as shown in FIG. 5 inexaggerated form, is given the highest incidence 26 relative to centerbody 25 because it doesn't have the benefit of the upwash from the othersurfaces and also has to operate at a much higher wing loading relativeto the large chord centerbody 25 in order to achieve an ellipticaldistribution of lift across the wingspan of the vehicle. Depending onthe design the aft swept wing 29 or 30 may have a negative incidence 28relative to the center body 25. The aft swept wing 29 or 30 has thebenefit of flying in the upwash of both the center body 25 and theforward swept wing 27/27A so depending on the particular design it mightneed a lower incidence to create the same amount of lift. Howevercountering this, like the forward swept wing 27, the aft swept wing 29or 30 also has a much smaller chord than the center of the center body25 and thus must have a much higher wing loading in order to achieve anelliptical distribution of lift across the entire vehicle.

It also can be beneficial for the center body 25 airfoil to be a reflexairfoil in order to provide a positive vehicle pitch-up moment fortrimming the vehicle with a farther forward center of gravity forincreased stability. Because of the large chord of this center body 25,this can be achieved without risk of overloading and stalling theleading edge. The aft swept wing 30 shown in FIG. 3 will generally needmore positive incidence than will the aft swept wing 29 arrangementshown in FIG. 2 because of its more highly swept design. Balancing ofthe lift between the forward and aft swept wings with minimal controlsurface deflections on the configuration shown in FIG. 3 can be moreeasily achieved with just curvi-linear dihedral as described below thanit can for the configuration shown in FIG. 2 which is more likely torequire different incidence angles between the forward and aft sweptwings 27 and 29.

D. How to Establish the Wing Dihedral

Once the wing incidence for low speed flight has been established thatachieves the closest thing to an elliptical distribution of lift acrossthe wingspan with the control surfaces in a neutral position, theconfiguration can be swept to the highest sweep position and wingcurvi-linear dihedral can be used to again achieve an ellipticaldistribution of lift. This is done as follows; At this high sweep angle,generally the farther forward on the vehicle the more the wing needs tobe increased in angle of attack in order to generate sufficient lift.Also, the farther aft the more the wing needs to be reduced in angle ofattack because this section of the wing is flying in the upwash of allthe wing sections in front. This can be achieved by having the dihedralsimilar to that shown in FIG. 4 which is a view of the aircraft in FIG.2 looking aft perpendicular to the pitch axis 39. Looking to the left onthe frontal view of the aircraft in FIG. 4, the center body 25 and theforward swept wing 27 slowly curves up with more and more dihedral. Whenthe wing is swept relative to the direction of travel, the increasingdihedral angle the farther outboard/forward effectively providesincreasing angle of attack at that point on the airfoil to compensatefor the effect of wing sweep. Looking to the right, the center body 25and the aft swept wing 29 generally also curve up more and more in acurvi-linear dihedral. When the wing is swept, because of the sweptangle of the aft swept wing 29, increasing dihedral angle results in areduction in the angle of attack of the local airfoil to compensate forthe increased relative lift that would otherwise occur due to the upwashfrom the forward 27 and center 25 wing sections. The larger sweepbackangle of the quarter chord 1 on the aft facing transition area 13 andaft facing wing 30 can also act to reduce the roll effect associatedwith increasing vehicle sweep angles and thus reduce the required amountof dihedral on these portions of the aircraft. Use of dihedral couldalso be used to solve local upwash problems if found necessary ordesirable. For example if during high speed for a particular design thetransition from the highly swept leading edge 17 to the more modestsweep of leading edge 31 for aft swept wing 29 for the configurationshown in FIG. 2 caused a locally high upwash and thus higher thandesired angle of attack at the wing root of wing 29, the dihedral couldbe increased significantly in this area where the wing 29 meets thetransition area 13 as shown in exaggerated form in FIG. 4 at point 33.If the upwash effect of this transition area is less outboard from thispoint the dihedral angle might drop as shown in exaggerated form atpoint 50 in FIG. 4 in order to maintain the proper distribution of liftalong the span. Although this is not anticipated to be a problem, if itwere a problem then leading edge flaps might be located at point 33 onthe root of wing 29 to prevent stall from the potential high angle ofattack air coming off the highly swept leading edge 17 of the transitionarea 13 during high g maneuvers.

As the vehicle pulls positive g′s while maneuvering, the wings can tendto bend up to higher angles effectively creating greater dihedral anglessince the vehicle isn't necessarily a pure spanloader where the weightis distributed exactly where the lift is. Theoretically if the wingstiffness and mass distribution of the vehicle is just right, the wingdeflections acting like increased dihedral can compensate for thegreater upwash airflows generated by the maneuvering and thus smallercontrol surface movements would be required to retrim the aircraft at aparticular design airspeed.

E. How to Achieve Yaw Stability and Control

In the embodiment of FIG. 2, a vertical fin 23 provides yaw stabilityfor the vehicle at the highly swept angle used in high speed flightwhere the vehicle is flying in direction 3. When a vertical fin 23 isused for yaw stability and control it rotates about a primarily verticalaxis and remains pointing generally in the direction of flight. Becausethe aft swept wing 29 or 30 is flying in the upwash from the center body25 and forward swept wing 27/27A, wing 29 or 30 tends to experience lessdrag than the forward swept wing 27/27A. On previous Oblique wingaircraft this caused a significant yaw moment that was difficult tohandle. The embodiment of the current invention shown in FIG. 2 has lessof a problem in that regard in that the chord line 1 of the forwardswept wing 27 is now forward of the chord line of the aft swept wing 29so that the inboard wing forces associated with wing dihedral, spaced asignificant distance apart fore and aft, generates a yawing moment inthe opposite direction to partially or fully eliminate this effect. Theconfiguration shown in FIG. 3 with aft swept wing 30 can also tend tolimit this yawing effect because the increased upwash on the aft sweptwing 30 relative to the forward swept wing 27A can be partly or fullycompensated by the larger swept back angle of the wing 30 versus wing 29and thus lower aspect ratio.

Another way to counter the traditional yawing effect of an oblique wingaircraft is to place the center of thrust 52 of the engines behind thevehicle center of gravity on the center body 25 as shown in FIG. 2. Withthis approach, as the vehicle yaws counter-clockwise for high speedflight, the engine thrust line 37 ends up moving to the right relativeto the vehicle center of gravity near 25 as shown in FIG. 2. With theengine exhaust exiting the aircraft behind the vehicle center of gravityat point 52, thrust deflectors can also be used to generate yawingmoments. For configurations where the engines are mounted in a fixedposition in the center body 25, thrust deflectors are needed anyway toadjust for the changing sweep/yaw angles of the vehicle used from low tohigh speed flight and keep the engine exhaust pointed primarily aft.Another approach is of course to have two engines spaced laterally fromeach other and to vary the thrust between the two engines to counter theyaw forces. In high speed, a vertical fin 23 mounted on the aft sweptwing 29 or 30 provides a lot of tail volume and a powerful yaw controlsystem. If the wing still tends to want to yaw to lower sweep angles, itis advantageous to locate more, if not all, of the vertical fin 23 abovethe wing 29 because if it is generating an inward force countering thisyawing effect, it will also act as a winglet to improve the efficiencyof the aft swept wing 29 or 30. Having the vertical fin above the wing29 also will of course help with ground clearance on landing. A wingletto be most effective needs to generate an inboard force for the arealocated above the wing and an outboard force for the area located below.Proper balancing of the area above and below the wing is desirable tomake the vertical tail 23 operate the most efficiently as a winglet. Themore the wings want to unsweep, the more vertical tail 23 area isdesired above the wing 29 and vice versa.

It should be noted that if a line describing the centerline of aninboard oriented force from the vertical fin 23 passes over the pitchaxis 39 of the vehicle, the vertical fin 23 will be generating a vehiclepitch up moment. The reverse occurs if the centerline of vertical finforces passes under the pitch axis of the vehicle or for fin forces inthe opposite direction. This is a further coupling of axes which issignificant when the vehicle is at a highly swept angles for high speedflight. To decouple vertical fin 23 forces from vehicle pitching momentsas much as possible, the vertical fin 23 can be canted outboard so thatthe vertical fin 23 generates forces pointed more closely to a linepassing through the vehicle pitch axis when the vehicle is at high sweepangles.

At low speed and low aircraft sweep angles, the vertical fin 23 may notbe located far enough behind the aircraft center of gravity to provideall the necessary yaw control by generating side forces alone. Underthese conditions the aircraft can use drag devices like ailerons thatsplit into an upper and lower segment like the B-2 to increase drag onone side or the other of the aircraft. The vertical fin 23 incombination with an adjacent aileron 35 can also be positioned to createdrag such as rotating the fin 23 counter-clockwise from its positionshown in FIG. 2. This will cause the fin 23 to stop operating as awinglet thus increasing wing drag and reducing wing lift. The aileron 35could then be deflected trailing edge down to compensate for the reducedlift and to generate more drag as well.

In this low sweep position the vertical fin 23 can also be helpful inproviding a direct side force capability to make it easier for the pilotto maneuver onto a runway without having to bank the aircraft andpotentially allowing a shorter landing gear with less fear of wingtipstrikes on the runway.

F. How to Achieve Pitch Control

Pitch for this aircraft is defined as rotation about axis 39 as shown in

FIGS. 2 and 3. The configuration shown in FIG. 3 uses a relativelystraightforward elevon control system. However because of the unusualdefinition of the pitch axis 39, the left elevon 61 of wing 30 providescloser to pure pitch control and right elevon 40 provides closer to pureroll control.

The configuration shown in FIG. 2 is more challenging and unusual.Trailing edge flaps 41 and 43 can be used to pitch the aircraft in thesame manner used on previous Oblique flying Wing aircraft. Deflectingthe trailing edges up causes a reduction in lift behind the vehiclecenter of gravity causing the vehicle to pitch nose up and vice versa.In addition, if more pitch power is required, deflecting flaps 36 andaileron 35 trailing edge up will cause more nose up pitching moment. Inorder to counter the left roll that this normally would cause, aileron40 (and potentially also flaps 38 and 42) could also be deflectedtrailing edge up. Flap 42 has little or no effect on pitch but does helpto counter the roll. Although the forces generated by flaps 38 andaileron 40 are actually slightly ahead of the pitch axis 39 through thecenter of gravity of the vehicle and thus actually generate a nose downmoment, it is much less powerful than the pitch up moment from surfaces35 and 36. A pure nose up pitching moment with these control surfacescan of course be generated in other ways such as inboard flap 36deflected trailing edge up and aileron 35 deflected modestly theopposite direction to cancel out the roll effect.

Pitch control can also be augmented in other ways such as engine thrustvectoring. Deflecting engine exhaust up behind the center of gravitywill generate a fuselage pitch up moment. Engine exhaust blowing overthe top or bottom of trailing edge flap 43 and 41 can increase theeffectiveness of these flaps. If more pitch stability and control poweris desired, flap 43 could be enlarged by extending it out to the rightso that the right end of flap 43 is cantilevered out to the right of theline of the trailing edge 19. The center of pressure effect of thisexpanded flap area to the right of the trailing edge 19 is far aft ofthe vehicle pitch axis 39 and center of gravity and because its to theright of the vehicle center it also acts to counter the roll effect offlaps 36 when used for pitch control. More pitch stability can beachieved by moving the center of gravity of the vehicle farther forwardand using a reflex airfoil on the center body 25 in order to retrim thevehicle with enough nose up pitching moment and still carry enough liftwith the aft swept wing 29 to get the optimum elliptical distribution oflift across the full wingspan of the aircraft. In general, for thepresent invention, a reflex airfoil on the centerbody 25 is preferred.

When the aircraft is flying at a swept angle, the vertical fin 23 canalso generate pitching moments about the pitch axis. Normally, for mostaircraft, coupling between controls is not considered desirable and thepitch effect of the vertical fin 23 (if located only above or below thewing 29) can be reduced by canting it outboard so its force whichotherwise would tend to produce a nose up or down moment respectively iscancelled out by its lifting force behind the center of gravity whichcauses an opposing nose down or up moment respectively. However byhaving an upper and lower rudder the aircraft could move the ruddersdifferentially to create vehicle pitching moments if that were desiredwhile neutralizing vehicle yawing moments.

G. How to Achieve Roll Control

The roll axis for this aircraft is defined as perpendicular to the pitchaxis 39. Roll control is achieved similar to other aircraft with the useof ailerons 35 and 40 which can be augmented if necessary by flaps 38and 36. Since surfaces 35 and 36 have more of a pitch effect on theaircraft than surfaces 38 and 40, flaps 41 and 43 could act opposite tosurfaces 35 and 36 to cancel out the pitch effect. Alternatively onlyaileron 40 might be used for small roll adjustments since it has littleeffect about the pitch axis 39.

H. Engine Integration

The engines can be located in pods under, or over, the wing that rotateto keep the engine pointed into the relative wind, or the engines can bebuilt into the center body 25. When the engines are built into the wing,the intakes and exhaust nozzles have to be able to operate at thedifferent aircraft yaw/sweep angles. Numerous thrust vectoring nozzleshave been developed such as those on the F-22 Raptor and the V/STOLversion of the Joint Strike

Fighter that could be utilized in a similar approach on this vehicle.Air intakes have also been built to efficiently take in air fromdifferent directions as would be required for engines built into thewing in this invention. The F-15 Eagle is such an example where theintake ramp rotates down approximately 45 degrees to intake airefficiently with the aircraft at different angles of attack. Simplerengine intake configurations for this application would be possiblesince the aircraft generally won't be flying supersonically at a lowsweep angle and generally won't be flying slow in a high sweep angle. Itis generally easier to integrate jet engines inside the wing of thisOblique Flying Wing than previous Oblique Wing Aircraft because of thegreater depth and length possible with the large chord center body 25.Otherwise the engine integration in this invention is similar to thoseproposed in earlier Oblique All Flying Wings.

I. Pitch and Roll Stability

FIG. 2 shows a dotted ¼ chord line 1 for this vehicle with aft sweptwing 29. The quarter chord line 1 of the forward swept wing 27 and itstransition area 15 are less distance in front of the pitch axis 39 thanthe quarter chord of the aft swept wing 29 and its transition area 13are behind the pitch axis 39. Also the aft swept structures 29 and 13are flying in the upwash from everything in front and they also have awinglet/vertical fin 23, so when a gust comes along they tend to gainmore lift than the forward swept structures 15 and 27. All of thesefactors contribute to the pitch stability of the vehicle. In a flight ofa demonstrator aircraft having ballast that could be moved forward oraft to vary the center of gravity position, the aircraft wassuccessfully flown with centers of gravity between and including 31% and39.5% chord on the centerbody. This is a remarkably far aft center ofgravity. However to further improve on the stability, the forward sweptwing 27 can be modified to pick up less lift with vehicle angle ofattack. For example, 45 degree composite plies in the top skin of thewing 27 and minus 45 degree plies in the bottom skin can cause the wing27 to twist leading edge down when the wing 27 is deflected up eitherfrom a gust or from the vehicle pitching up. Another way to achieve thesame result is to add a wingtip 51 as shown in FIG. 3, whose center ofpressure is behind the elastic axis 55 of the wing 27A, so when the winghits a gust or the vehicle pitches up, the wing will twist leading edgedown. This acts as a natural gust alleviation system and also increasesvehicle pitch stability. This wingtip 51, because it is unswept in highspeed flight, would be a very thin structure like an F-104 wing tominimize compressibility drag. Its other benefit is that it increasesthe wetted aspect ratio of the vehicle at high sweep angles byincreasing the wing span. A thin wing is heavier than an equivalentthick wing but because the wingtip 51 is a short structure withrelatively low bending loads the impact is minimal and achieves the samewingspan increase when the vehicle is at a highly swept angle of asignificantly longer in-line extension of the wing 27A which would havemore surface area and possibly more weight. In addition, by activelycontrolling an aileron surface 53 on this wingtip 51 the wing 27A can beactively twisted down to counter an up gust of air. Also having theaerodynamic center of the wingtip 51 behind the elastic axis 55 of thewing 27A provides damping and twist stability to the wing 27A. Theaileron 53 also can be actively driven to provide additional damping ofthe wing 27A twist to delay the onset of flutter and to offset thede-stabilizing coupling that the offset wing tip creates between torsionand bending modes of the wing. A sensor to detect air gusts in advanceand an accelerometer and rate gyro near the wingtip can be used in aclosed circuit control system to try to minimize vertical accelerationsat the wingtip and provide a damping force for torsional motion.Minimizing vertical accelerations at the wingtip should also help theride quality of the entire vehicle. For a wing 27A that isn't very stiffin torsion, the aileron 53 will always work counter to a conventionalaileron in that additional lift on the aileron 53 causes the wing totwist leading edge down resulting in a net loss instead of gain in lift.However the control surfaces 40 and 38 just inboard of the wingtip 51still works in a conventional manner like ailerons and the wingtipaileron 53 can be used to prevent twisting of the wing that usuallycauses control reversal for jet transports at high speed. It should beunderstood that the forward swept wing wingtip extension 51 shown inFIG. 3 could be used on the configuration shown in FIG. 2. It will alsobe understood that the more conventional forward swept wing design 27shown in FIG. 2 could also be used on the configuration shown in FIG. 3.

It will also be understood that the aircraft could have a system likethe B-2 bomber that can move fuel between forward and aft tanks toprecisely control the center of gravity of the vehicle at all but a zerofuel state.

J. How to Have a Stable Platform on the Ground Yet Still be Able toRotate for Takeoff

There are several ways to allow the aircraft to be very stable whilesitting on the ground on its landing gear yet be capable of rotatingeasily for takeoff. This vehicle already has better stability on theground than previous OAW (Oblique All Wing) aircraft since the vehicleis spread farther fore and aft relative to the pitch axis 39 allowing alonger potential wheel base and a greater pitching moment capability.The vehicle also has a smaller moment of inertia in the roll and yawaxis because of a smaller wingspan and the weight of the vehicle beingmore concentrated near the centerline.

The use of the vertical fin 23 for direct side force could be used sothe aircraft doesn't have to bank significantly on landing and as aresult can have a shorter landing gear. Also engine thrust could be usedto assist in pitching up the aircraft for takeoff either with a thrustline below the vehicle center of gravity or by deflecting the engineexhaust up behind the aft landing gear bogies. Another approach is tohave one or more landing gear struts situated well behind the vehiclecenter of gravity to provide good stability on the ground but which canbe fully or partially retracted during the takeoff roll, before theother landing gear struts, to allow easy rotation. A wide landing gearspacing is generally preferred if the runways and taxiways willaccommodate it for better stability on the ground and also so it is lesslikely that the aft swept wingtip will contact the ground especiallysince it may be desirable to have some of the vertical fin/winglet 23extend below the wing.

If the vehicle has all steerable landing gear bogies the aircraft couldtaxi in its lengthwise direction making it possible to get into tightspaces and allowing the aircraft to be densely parked next to othersimilar aircraft.

The current invention also has other potential advantages. At high speedthe aircraft could yaw to a low sweep angle in order to potentiallyachieve high maneuverability and high compressibility drag such asduring air to air combat to slow down and turn rapidly to force anadversary to overshoot its position. Also at low speed the aircraftcould yaw to a high sweep angle to increase induced drag such as duringan approach to landing to achieve a steeper glide slope angle. Theaircraft could also potentially handle cross wind landing better thanother aircraft.

Although an all-wing aircraft has been shown and is much the preferredembodiment of this invention, it should be stated that this invention isalso applicable to an oblique wing and fuselage configuration as well.In that case a fuselage is preferably mounted under the oblique wing ofthis invention similar to previous oblique wing/fuselage aircraft.

Numerous other variations in the aircraft of the invention, within thescope of the appended claims, will occur to those skilled in the art inlight of the foregoing disclosure. As various changes could be made inthe above constructions without departing from the scope of theinvention, it is intended that all matter contained in the abovedescription or shown in the accompanying drawings shall be interpretedas illustrative and not in a limiting sense.

1. An all-wing oblique wing aircraft designed for reduced surface areato volume ratio, said aircraft having an oblique wing comprising: aforward swept wing segment on one side of the wing and an aft swept wingsegment on the opposite side of the wing and characterized by furthercomprising a center oblique airfoil section connecting said forward andaft swept wing segments, said center oblique airfoil section having alarger chord length near its centerline than the chord length of eitherof said forward or aft swept wing segments; the chord length of saidcenter oblique airfoil section tapering down more rapidly than saidforward or aft wing segments as said center oblique airfoil sectionextends outboard toward said forward and aft swept wings; said centeroblique airfoil section not being shaped solely to function as acircular fairing to fill the gap between an oblique wing and a fuselageat different oblique wing angles; said center airfoil section not beinga second wing in an X wing configuration.
 2. The oblique wing aircraftof claim 1, where the leading edge of said aft swept wing segment isswept less than the leading edge of the aft side of said center obliqueairfoil section.
 3. The oblique wing aircraft of claim 1, where theforward swept wing segment has a higher incidence than the centerairfoil section and aft swept wing segment.
 4. The oblique wing aircraftof claim 1, where all or part of the half chord line of said centeroblique airfoil section is more highly swept than the half chord line ofthe forward swept wing segment.
 5. The oblique wing aircraft of claim 4,where all or part of said half chord line of said center oblique airfoilsection is swept more than 10 degrees more than the half chord line ofthe forward swept wing segment.
 6. The oblique wing aircraft of claim 4,where all or part 5 of said half chord line of said center obliqueairfoil section is swept more than 15 degrees more than the half chordline of the forward swept wing segment.
 7. The oblique wing aircraft ofclaim 4, where all or part of said half chord line of said centeroblique airfoil section is swept more than 20 degrees more than the halfchord line of the forward swept wing segment.
 8. The oblique wingaircraft of claim 1, where over 25% of the total planform area can fitinside the largest circle that can be laid fully over the planform area.9. The oblique wing aircraft of claim 1, where over 33% of the totalplanform area can fit inside the largest circle that can be laid fullyover the planform area.
 10. The oblique wing aircraft of claim 1, whereover 40% of the total planform area can fit inside the largest circlethat can be laid fully over the planform area.
 11. The oblique wingaircraft of claim 1, where over 50% of the total planform area can fitinside the largest circle that can be laid fully over the planform area.12. The oblique wing aircraft of claim 1, where over 50% of the outlineof the wing in planform is composed of essentially two sets of parallellines.
 13. The oblique wing aircraft of claim 1, where the leading edgeof said aft swept wing segment is swept less than the leading edge ofthe aft transition area of said center oblique airfoil section.
 14. Theoblique wing aircraft of claim 13, where a trailing edge flap attachedto the aft transition area on the center oblique airfoil section extendsout in the direction of the forward swept wing segment and aft relativeto a line described by the trailing edge of the forward transition area.15. The oblique wing aircraft of claim 1, where laminar flow airfoilsare used on forward or aft swept wing segments.
 16. The oblique wingaircraft of claim 1 wherein a propulsion system for the aircraft ismounted inside the center airfoil section.
 17. The oblique wing aircraftof claim 16 wherein the propulsion system comprises two jet engines.